Showing papers on "Leading edge published in 2005"

Bow Shock and Radio Halo in the Merging Cluster A520

Abstract: Chandra observations of the merging galaxy cluster A520 reveal a prominent bow shock with M = 21 This is only the second clear example of a substantially supersonic merger shock front in clusters Comparison of the X-ray image with that of the previously known radio halo reveals a coincidence of the leading edge of the halo with the bow shock, offering an interesting experimental setup for determining the role of shocks in the radio halo generation The halo in A520 apparently consists of two spatially distinct parts, the main turbulence-driven component and a cap-like forward structure related to the shock, where the latter may provide preenergized electrons for subsequent turbulent reacceleration The radio edge may be caused by electron acceleration by the shock If so, the synchrotron spectrum should have a slope of α 12 right behind the edge, with quick steepening farther away from the edge Alternatively, if shocks are inefficient accelerators, the radio edge may be explained by an increase in the magnetic field and density of preexisting relativistic electrons due to gas compression In the latter model, there should be radio emission in front of the shock with the same spectrum as that behind it, but 10-20 times fainter If future sensitive radio measurements do not find such preshock emission, then the electrons are indeed accelerated (or reaccelerated) by the shock, and one will be able to determine its acceleration efficiency We also propose a method to estimate the magnetic field strength behind the shock, based on measuring the dependence of the radio spectral slope upon the distance from the shock In addition, the radio edge provides a way to constrain the diffusion speed of the relativistic electrons

A rod-airfoil experiment as a benchmark for broadband noise modeling

Abstract: A low Mach number rod-airfoil experiment is shown to be a good benchmark for numerical and theoretical broadband noise modeling. The benchmarking approach is applied to a sound computation from a 2D unsteady-Reynolds-averaged Navier–Stokes (U-RANS) flow field, where 3D effects are partially compensated for by a spanwise statistical model and by a 3D large eddy simulation. The experiment was conducted in the large anechoic wind tunnel of the Ecole Centrale de Lyon. Measurements taken included particle image velocity (PIV) around the airfoil, single hot wire, wall pressure coherence, and far field pressure. These measurements highlight the strong 3D effects responsible for spectral broadening around the rod vortex shedding frequency in the subcritical regime, and the dominance of the noise generated around the airfoil leading edge. The benchmarking approach is illustrated by two examples: In both cases, the ability of computational fluid dynamics to model the source mechanisms and of the CAA approach to predict the far field are assessed separately.

Self aligned wrap around shield for perpendicular magnetic recording

Abstract: A write element for use in perpendicular magnetic recording. The write element including a write pole and a self aligned wrap around shield that can have a trailing shield gap thickness that is different from its side shield gap thickness. The materials making up the trailing shield gap and the side shield gaps can be different materials or can be the same material deposited in two different steps. The side or wrap around portions of the trailing shield can extend down to the level of the leading edge of the write pole or can terminate at some point between the levels of the leading and trailing edge to form a partial wrap around trailing shield.

Investigation of Three-Dimensional Dynamic Stall Using Computational Fluid Dynamics

Abstract: Numerical simulation of three-dimensional dynamic stall has been undertaken using computational fluid dynamics. The full Navier–Stokes equations, coupled with a two-equation turbulence model, where appropriate, have been solved on multiblock strucured grids in a time-accurate fashion. Results have neen obtained for wings of square planform and of NACA 0012 section. Efforts have been devoted to the accurate modeling of the flow near the wing tips, which, for this case, were sharp without tip caps. The obtained results revealed the time evolution of the dynamic stall vortex, which, for this case, takes the shape of a capital omega+spanning the wing. The obtained results compare well against experimental data both for the surface pressure distribution on the wing and the flow topology. Of significant importance is the interaction between the three-dimensional dynamic stall vortex and the tip vortex. The present results indicate that once the two vortices are formed both appear to originate from the same region, which is located near the leading edge of the tip. During the ramping of the wing, the two vortices grow significantly in size. The dynamic stall vortex dettaches from the wing in the inboard region but remains close to the wing’s leading edge near the tip. The overall configuration of the developed vortical system takes a form. To our knowledge, this is the first detailed numerical study of three-dimensional dynamic stall appearing in the literature.

Optimization of Flapping Airfoils For Maximum Thrust and Propulsive Efficiency

Abstract: The thrust and/or propulsive efficiency of a single flapping airfoil is maximized by using a numerical optimization method based on the steepest ascent. The flapping motion of the airfoil is described by a combined sinusoidal plunge and pitching motion. Optimization parameters are taken to be the amplitudes of the plunge and pitching motions and the phase shift between them at a fixed flapping frequency. Two-dimensional, unsteady, low-speed, laminar, and turbulent flows are computed by using a Navier‐Stokes solver on moving overset grids. Computations are performed in parallel in a computer cluster. The optimization data show that high thrust values may be obtained at the expense of propulsive efficiency. For a high propulsive efficiency, the effective angle of attack of the airfoil is reduced, and large-scale vortex formations at the leading edge are prevented.

Effect of Airfoil Aerodynamic Loading on Trailing Edge Noise Sources

Abstract: A previous experimental investigation of the broadband self noise radiated by an industrial cambered controlled-diffusion airfoil embedded in an homogeneous flow at low Mach number has been extended to various aerodynamic loadings. The instrumented airfoil is placed at the exit of an open-jet anechoic wind tunnel, with a jet width of about four chord lengths. Sound is measured in the far field at the same time as the statistical properties of the wall-pressure fluctuations close to the trailing edge. A new set of mean wall-pressure data has been collected on this airfoil at a chord Reynolds number of 2.9 x 105, which provides some insight on the Reynolds-number effect. Two previously investigated flow regimes with different statistical behaviors are investigated by changing the angle of attack from 8 to 15 deg. They respectively correspond to the nearly separated boundary layer with vortex shedding at the trailing edge and to the turbulent boundary layer initiated by a leading-edge separation.

Three-dimensional transition in the wake of bluff elongated cylinders

Abstract: Despite little supporting evidence, there appears to be an implicit assumption that the wakes of two-dimensional bluff bodies undergo transition to three-dimensional flow and eventually turbulence, through the same sequence of transitions as observed for a circular cylinder wake. Previous studies of a square cylinder wake support this assumption. In this paper, the transition to three-dimensional wake flow is examined for an elongated cylinder with an aerodynamic leading edge and square trailing edge. The three-dimensional instability modes are determined as a function of aspect ratio (, is more unstable than Mode A. These results suggest that the transition scenario for elongated bluff bodies may be distinctly different to short bodies such as circular or square cylinders. At the very least, the dominant spanwise wavelength in the turbulent wake is likely to be much longer than that for a circular cylinder wake. In addition, the reversal of the ordering of occurrence of the two modes with the different spatial symmetries is likely to affect the development of spatio-temporal chaos as a precursor to fully turbulent flow.In conjunction with prior work, the current results indicate that nearly all three-dimensional instabilities of the vortex street can be identified as one of only a handful of transition modes.

Infall of the Elliptical Galaxy NGC 1404 into the Fornax Cluster

Abstract: We use three Chandra observations, totaling 134.3 ks, to constrain the dynamical motion of NGC 1404 falling toward the dominant elliptical NGC 1399 through the Fornax Cluster gas. The surface brightness profile of NGC 1404 shows a sharp edge at ~8 kpc from its center in the direction of NGC 1399, characteristic of jumplike temperature and density discontinuities from ram pressure stripping of the galaxy gas, caused by its motion through the surrounding intracluster medium (ICM). We find that the temperature of the galaxy gas inside the edge is ~2.8 times cooler (kT = 0.55 keV with abundance A = 0.73 Z☉) than the cluster gas (kT = 1.53 keV, A = 0.42 Z☉). We use the shape of the surface brightness profile across the edge to fit the position of the edge, the power-law behavior of NGC 1404's density distribution in the leading direction, and the density discontinuity at the edge. The electron density inside the edge (3.9-4.3) × 10-3 cm-3 depends strongly on the gas abundance, while the density of the ICM (7-8) × 10-4 cm-3 depends strongly on the assumed geometry (relative distance) between NGC 1404 and NGC 1399. The corresponding pressure jump of 1.7-2.1 across the leading edge of the galaxy and the cluster free-stream region implies near sonic motion (Mach number 0.83-1.03) for NGC 1404 with a velocity 531-657 km s-1 relative to the surrounding cluster gas. The inclination angle of the motion, inferred using the relative radial velocity between NGC 1404 and 1399 as representative of that between NGC 1404 and the cluster ICM, is uncomfortably large (40°) given the sharpness of the surface brightness edge, suggesting either a nonzero impact parameter between NGC 1404 and 1399 or that NGC 1399 is also moving radially with respect to the cluster ICM.

Dynamics of turbulence spreading in magnetically confined plasmas

Abstract: A dynamical theory of turbulence spreading and nonlocal interaction phenomena is presented. The basic model is derived using Fokker–Planck theory, and supported by wave-kinetic and K-ϵ type closures. In the absence of local growth, the model predicts subdiffusive spreading of turbulence. With local growth and saturation via nonlinear damping, ballistic propagation of turbulence intensity fronts is possible. The time asymptotic front speed is set by the geometric mean of local growth and turbulent diffusion. The leading edge of the front progresses as the turbulence comes to local saturation. Studies indicate that turbulence can jump gaps in the local growth rate profile and can penetrate locally marginal or stable regions. In particular, significant fluctuation energy from a turbulent edge can easily spread into the marginally stable core, thus creating an intermediate zone of strong turbulence. This suggests that the traditional distinction between core and edge should be reconsidered.

A numerical study on the formation of diffusion flame islands in a turbulent hydrogen jet lifted flame

Abstract: This paper presents a numerical study on the formation of diffusion flame islands in a hydrogen jet lifted flame. A real size hydrogen jet lifted flame is numerically simulated by the DNS approach over a period of about 0.5 ms. The diameter of hydrogen injector is 2 mm, and the injection velocity is 680 m/s. The lifted flame is composed of a stable leading edge flame, a vigorously turbulent inner rich premixed flame, and a number of outer diffusion flame islands. The relatively long-term observation makes it possible to understand in detail the time-dependent flame behavior in rather large time scales, which are as large as the time scale of the leading edge flame unsteadiness. From the observation, the following three findings are obtained concerning the formation of diffusion flame islands. (1) A thin oxygen diffusion layer is developed along the outer boundary of the lifted flame, where the diffusion flame islands burn in a rather flat shape. (2) When a diffusion flame island comes into contact with the fluctuating inner rich premixed flame, combustion is intensified due to an increase in the hydrogen supply by molecular diffusion. This process also works for the production of the diffusion flame islands in the oxygen diffusion layer. (3) When a large unburned gas volume penetrates into the leading edge flame, the structure of the leading edge flame changes. In this transformation process, a diffusion flame island comes near the leading edge flame. The local deficiency of oxygen plays an important role in this production process.

Quantification of microtubule nucleation, growth and dynamics in wound-edge cells.

Abstract: Mammalian cells develop a polarized morphology and migrate directionally into a wound in a monolayer culture. To understand how microtubules contribute to these processes, we used GFP-tubulin to measure dynamic instability and GFP-EB1, a protein that marks microtubule plus-ends, to measure microtubule growth events at the centrosome and cell periphery. Growth events at the centrosome, or nucleation, do not show directional bias, but are equivalent toward and away from the wound. Cells with two centrosomes nucleated approximately twice as many microtubules/minute as cells with one centrosome. The average number of growing microtubules per μm 2 at the cell periphery is similar for leading and trailing edges and for cells containing one or two centrosomes. In contrast to microtubule growth, measurement of the parameters of microtubule dynamic instability demonstrate that microtubules in the trailing edge are more dynamic than those in the leading edge. Inhibition of Rho with C3 transferase had no detectable effect on microtubule dynamics in the leading edge, but stimulated microtubule turnover in the trailing edge. Our data demonstrate that in wound-edge cells, microtubule nucleation is non-polarized, in contrast to microtubule dynamic instability, which is highly polarized, and that factors in addition to Rho contribute to microtubule stabilization.

Prediction of Turbulent Inflow and Trailing-Edge Noise for Wind Turbines

Abstract: Two dominant sources of aeroacoustic noise for operating wind turbines are thought to be noise due to the interaction of atmospheric turbulence with the leading edge of the blades and turbulent boundary layer noise from the trailing edges of the blades. This study examines two new models for these noise sources: one is a simplified version of a previously implemented model of turbulent inflow noise and another is a more sophisticated version of a turbulent boundary layer trailing-edge noise model. Comparisons between simplified and sophisticated models are made as well as comparisons to measured data. The previously developed model for inflow turbulence noise is able to predict the differences in sound level between different airfoil shapes with good accuracy. The model is based on the acoustic analogy and unfortunately is also computationally intensive, making it prohibitively expensive for wind turbine design. A simplified version of this model, which is based on a geometrical analysis of the airfoils, is presented. It is based on the observation that the difference in inflow turbulence noise between an airfoil and a flat plate can be described by a straight line, with the only free parameter being the slope of this line. The simplified model relates this slope to the relative thickness of the airfoil at two points along the chord. Results are presented, which show that such a simple model can reproduce the results of the more sophisticated model with good accuracy below a Strouhal number based on chord of approximately 75. A new turbulent boundary layer trailing-edge noise model was also developed. This model is more complex than the currently implemented semi-empirical algoritms and more accurately models the physical processes that create boundary layer noise. The model uses boundary layer parameters to estimate the trailing edge noise on both sides of a given airfoil. The sound pressure levels for four different airfoils are calculated using this new method and compared to experimental data. The results from this method are also compared to the empirical relations of the original model based on the work of Brooks, Pope, and Marcolini. The accuracy of the new method is as good or slightly better than the original empirical algorithms for most cases studied. It is unclear whether the new trailing- edge noise model is sensitive enough to predict noise differences between airfoil shapes, but further refinement may increase its accuracy.

Delta wing stall and roll control using segmented piezoelectric fluidic actuators

Abstract: The separated flow around a balance-mounted, 60-deg sweptback, semispan delta wing with a sharp leading edge was controlled using zero-mass-flux periodic excitation from a segmented leading-edge slot. Excitation was generated by cavity-installed piezoelectric actuators operating at resonance with amplitude modulation (AM) and burst mode (BM) signals being used to achieve reduced frequencies (scaled with the freestream velocity and the root chord) in the range from O(1) to O(10). Results of a parametric investigation, studying the effects of AM frequency, BM duty cycle and frequency, excitation amplitude, location of the actuation along the leading edge, and optimal phase difference between the actuators, as well as the Reynolds number, are reported and discussed

The Dynamics of the Horseshoe Vortex and Associated Endwall Heat Transfer—Part I: Temporal Behavior

Abstract: Instantaneous flow topology and the associated endwall heat transfer in the leading-edge endwall region of a symmetric airfoil are presented. An experimental technique was employed that allowed the simultaneous recording of instantaneous particle image velocimetry flow field and thermochromic liquid-crystal-based endwall heat transfer data. The endwall flow is dominated by a horseshoe vortex that forms from reorganized impinging boundary layer vorticity. A corner vortex is shown to be a steady feature of the corner region, while a secondary vortex develops sporadically immediately upstream of the horseshoe vortex. The region upstream of the horseshoe vortex is characterized by a bimodal switching of the near-wall reverse flow, which results in quasi-periodic eruptions of the secondary vortex. The bimodal switching of the reverse flow in the vicinity of the secondary vortex is linked to the temporal behavior of the down-wash fluid on the leading edge of the foil. Frequency analysis of the flow field and endwall heat transfer data, taken together, indicate that the eruptive behavior associated with the horseshoe vortex occurs at a frequency that is essentially the same as the measured turbulence bursting period of the impinging turbulent endwall boundary layer.Copyright © 2005 by ASME

Asymmetric vortex shedding flow past an inclined flat plate at high incidence

Abstract: This paper reports an experimental investigation of the vortex shedding wake behind a long flat plate inclined at a small angle of attack to a main flow stream. Detailed velocity fields are obtained with particle-image velocimetry (PIV) at successive phases in a vortex shedding cycle at three angles of attack, α=20°, 25° and 30°, at a Reynolds number Re≈5,300. Coherent patterns and dynamics of the vortices in the wake are revealed by the phase-averaged PIV vectors and derived turbulent properties. A vortex street pattern comprising a train of leading edge vortices alternating with a train of trailing edge vortices is found in the wake. The trailing edge vortex is shed directly from the sharp trailing edge while there are evidences that the formation and shedding of the leading edge vortex involve a more complicated mechanism. The leading edge vortex seems to be shed into the wake from an axial location near the trailing edge. After shedding, the vortices are convected downstream in the wake with a convection speed roughly equal to 0.8 the free-stream velocity. On reaching the same axial location, the trailing edge vortex, as compared to the leading edge vortex, is found to possess a higher peak vorticity level at its centre and induce more intense fluid circulation and Reynolds stresses production around it. It is found that the results at the three angles of attack can be collapsed into similar trends by using the projected plate width as the characteristic length of the flow.

Application of Parallel Adjoint-Based Error Estimation and Anisotropic Grid Adaptation for Three-Dimensional Aerospace Configurations

Abstract: This paper demonstrates the extension of error estimation and adaptation methods to parallel computations enabling larger, more realistic aerospace applications and the quantification of discretization errors for complex 3-D solutions. Results were shown for an inviscid sonic-boom prediction about a double-cone configuration and a wing/body segmented leading edge (SLE) configuration where the output function of the adjoint was pressure integrated over a part of the cylinder in the near field. After multiple cycles of error estimation and surface/field adaptation, a significant improvement in the inviscid solution for the sonic boom signature of the double cone was observed. Although the double-cone adaptation was initiated from a very coarse mesh, the near-field pressure signature from the final adapted mesh compared very well with the wind-tunnel data which illustrates that the adjoint-based error estimation and adaptation process requires no a priori refinement of the mesh. Similarly, the near-field pressure signature for the SLE wing/body sonic boom configuration showed a significant improvement from the initial coarse mesh to the final adapted mesh in comparison with the wind tunnel results. Error estimation and field adaptation results were also presented for the viscous transonic drag prediction of the DLR-F6 wing/body configuration, and results were compared to a series of globally refined meshes. Two of these globally refined meshes were used as a starting point for the error estimation and field-adaptation process where the output function for the adjoint was the total drag. The field-adapted results showed an improvement in the prediction of the drag in comparison with the finest globally refined mesh and a reduction in the estimate of the remaining drag error. The adjoint-based adaptation parameter showed a need for increased resolution in the surface of the wing/body as well as a need for wake resolution downstream of the fuselage and wing trailing edge in order to achieve the requested drag tolerance. Although further adaptation was required to meet the requested tolerance, no further cycles were computed in order to avoid large discrepancies between the surface mesh spacing and the refined field spacing.

Airfoil with supplemental cooling channel adjacent leading edge

Abstract: An airfoil, and in a disclosed embodiment a rotor blade, has film cooling holes formed at a leading edge. A supplemental film cooling channel is positioned near the leading edge, but spaced toward the trailing edge from the leading edge. The supplemental film cooling channel directs film cooling air onto a suction wall. The supplemental film cooling channel air is generally directed to a location on the suction wall that has raised some challenges in the past. In a disclosed embodiment, the airfoil is provided with a thermal barrier coating, and the supplemental film cooling air protects this thermal barrier coating.

Wall configuration of axial-flow machine, and gas turbine engine

Abstract: A wall configuration of an axial-flow machine which can reduce the secondary flow loss is provided. A trough is formed between a blade and another blade in the blade row and extends in at least an axial direction of the blade row. The region where the trough is formed is axially between a leading edge and a trailing edge of the blade. A center line of the trough has a curvature in the same direction as a camber line of the blade. A maximum amplitude of the trough is located adjacent to an axial center of the blade or located axially between the axial center and the leading edge of the blade.

The Dynamics of the Horseshoe Vortex and Associated Endwall Heat Transfer—Part II: Time-Mean Results

Abstract: Time-mean endwall heat transfer and flow-field data in the endwall region are presented for a turbulent juncture flow formed with a symmetric bluff body. The experimental technique employed allowed the simultaneous recording of instantaneous particle image velocimetry flow field data, and thermochromic liquid-crystal-based endwall heat transfer data. The time-mean flow field on the symmetry plane is characterized by the presence of primary (horseshoe), secondary, tertiary, and corner vortices. On the symmetry plane the time-mean horseshoe vortex displays a bimodal vorticity distribution and a stable-focus streamline topology indicative of vortex stretching. Off the symmetry plane, the horseshoe vortex grows in scale, and ultimately experiences a bursting, or breakdown, upon experiencing an adverse pressure gradient. The time-mean endwall heat transfer is dominated by two bands of high heat transfer, which circumscribe the leading edge of the bluff body. The band of highest heat transfer occurs in the corner region of the juncture, reflecting a 350% increase over the impinging turbulent boundary layer. A secondary high heat-transfer band develops upstream of the primary band, reflecting a 250% heat transfer increase, and is characterized by high levels of fluctuating heat load. The mean upstream position of the horseshoe vortex is coincident with a region of relatively low heat transfer that separates the two bands of high heat transfer.

Timing estimation in an OFDM receiver

Abstract: A timing estimation system and methodology are provided. In particular, a first pilot is employed in conjunction with three acquisition stages. In the first stage, an attempt is made to observe the leading edge of the correlation curve associated with the first pilot symbol. In the second stage, a determination is made to confirm a leading edge was detected in the first stage by attempting to observe a trailing edge of the correlation curve. Furthermore, during this second stage, a frequency loop is updated to account for frequency offset. The third stage is for observing the trailing edge of the curve if it was not already observed in stage two. Upon detection of receipt of the first pilot, a second pilot can subsequently be employed to acquire fine symbol timing.

Mechanism of combustion dynamics in a backward-facing step stabilized premixed flame

Abstract: Combustion dynamics leading to thermoacoustic instability in a rearward-facing step stabilized premixed flame is experimentally examined with the objective of investigating the fluid dynamic mechanism that drives heat release rate fluctuations, and how it couples with the acoustic field. The field is probed visually, using linear photodiode arrays that capture the spatiotemporal distribution of CH* and OH*; an equivalence ratio monitor; and a number of pressure sensors. Results show resonance between the acoustic quarter wave mode of the combustion tunnel and a fluid dynamic mode of the wake. Under unstable conditions, the flame is convoluted around a large vortex that extends several step heights downstream. During a typical cycle, while the velocity is decreasing, the vortex grows, and the flame extends downstream around its outer edge. As the velocity reaches its minimum, becoming mostly negative, the vortex reaches its maximum size, and the flame collides with the upper wall; its leading edge folds, trapping reactants pockets, and its trailing edge propagates far upstream of the step. In the next phase, while the velocity is increasing, the heat release grows rapidly as trapped reactant’ pockets are consumed by flames converging towards their centers, and the upstream flame is dislodged back downstream. The heat release rate reaches its maximum halfway into the velocity rise period, leading the maximum velocity by about 90°. In this quarter-wave mode, the pressure leads the velocity by 90° as well, that is, it is in phase with the heat release rate. Numerical modeling results support this mechanism. Equivalence ratio contribution to the instability mechanism is shown to be minor, i.e., heat release dynamics are governed by the cyclical formation of the wake vortex and its interaction with the flame.

Turbine nozzle with trailing edge convection and film cooling

Abstract: An airfoil for a gas turbine engine has opposed pressure and suction sidewalls extending between a leading edge and a trailing edge. The airfoil includes an array of radially-spaced apart longitudinally-extending lands which define a plurality of trailing edge slots therebetween. Each of the trailing edge slots has an inlet in fluid communication with an interior of the airfoil and an exit in fluid communication with the trailing edge. At least one of the lands is tapered such that a width of the land measured in a radial direction decreases from the suction sidewall to the pressure sidewall.

Separation-Bubble-Transition Measurements on a Low-Re Airfoil Using Particle Image Velocimetry

Abstract: This paper presents experimental results on separation-bubble transition at low Reynolds number and low freestream turbulence, measured on an airfoil using particle image velocimetry (PIV). The two-dimensional PIV measurements have been performed over the suction surface of a low-Reynolds-number airfoil in a water tow-tank facility. Reynolds numbers, based on airfoil chord length and towing speed, of 40,000 and 65,000 have been examined at various angles of incidence, providing a range of streamwise pressure distributions and transitional separation-bubble geometries. The types of bubbles observed range from a short and thick bubble with separation near the leading edge of the airfoil, to a long and thin bubble with separation far downstream of the suction peak. The PIV measurements facilitate visualization of the vortex dynamics associated with separation-bubble transition. The growth of instability waves within the separated shear layer and eventual breakdown into turbulence is documented through the instantaneous vector fields. For all cases examined, large-scale vortex shedding and multiple reverse-flow zones are observed in the reattachment region. A technique for estimating the location of transition onset based on statistical turbulence quantities is presented, and comparisons are made to existing transition models.Copyright © 2005 by ASME

Experimental Study on Aerodynamic Characteristics of Unsteady Wings at Low Reynolds Number

Abstract: Aerodynamic forces and moment acting on wings of AR = 6 with heaving and feathering oscillations in a wind tunnel were measured at a low Reynolds number less than 10 4 . Airfoils of the wings examined are not streamlined, but they are in various profiles such as a flat plate with and without sharp leading edge, circular arc, and corrugated airfoils. By analyzing the sinusoidal aerodynamic forces and moment, it was found that some differences among airfoils were remarkable in both the mean values and the first harmonic amplitude of aerodynamic coefficients. A large perpendicular force is obtained in some airfoils, but the thrust was almost canceled by the drag in this low-Reynolds-number range during heaving motion alone. To get the maximum thrust, the optimal phase shift of combined heaving and feathering motion was required.

Net Heat Flux Reduction and Overall Effectiveness for a Turbine Blade Leading Edge

Abstract: The external cooling performance of a film cooled turbine airfoil can be quantified as a net reduction in heat transfer relative to the turbine airfoil without film cooling. This quantification is generally accomplished by using measurements of the adiabatic effectiveness and the change in heat transfer coefficients (hf /h0 ) for the film cooled surface to determine the net heat flux reduction (Δqr ). Although measurement of Δqr for laboratory models give an indication of the ultimate film cooling performance, this does not show how much the surface temperature of the airfoil is reduced by film cooling. Measurement of scaled surface temperatures can be accomplished by using laboratory models constructed so that the Biot number is matched with that of the actual airfoil. These measurements provide a scaled temperature distribution on the airfoil that is referred to as the overall effectiveness, φ. For the current study, measurements of Δqr and φ have been made for a simulated turbine blade leading edge. The simulated leading edge incorporated shaped coolant holes, and had three rows of coolant holes. Improvements due to the shaped holes were determined by comparisons with previously measured round hole configurations. Spatially distributed hf /h0 show increases of 5% to 15% for M = 1.0 and 10% to 30% for M = 2.0. Results show that local variation in Δqr much greater than variation in φ, but laterally averaged Δqr distributions are reasonable predictors of the laterally averaged φ distributions.Copyright © 2005 by ASME

Suppression of Control Reversal Using Leading- and Trailing-Edge Control Surfaces

Abstract: Control reversal is the loss, due to the flexibility of the primary aerostructure, of aircraft maneuvering loads induced by control surfaces. In recent years, attention has been given to the suppression of reversal through the use of distributed control surfaces. The authors study reversal behavior for a wing section with full-span leading-edge and trailing-edge control surfaces. The essential equations are developed by examining static aeroelastic responses. Analysis and experiments are presented. Specific trailing-edge to leading-edge control commands are identified to optimize performance. Although reversal is not eliminated in the experiments, the addition of the leading edge is shown to improve performance substantially. The research also identifies the adverse consequences of actuator flexibility.

Wind turbine rotor assembly and blade having acoustic flap

Abstract: A wind turbine includes a rotor assembly having at least one blade with a blade body defining a leading edge and a trailing edge and adapted for movement in response to a wind flow over the body to produce electricity. A rigid acoustic flap extends outward from the trailing edge, and a distal end of the acoustic flap is substantially smooth and continuous. The flap reduces acoustic noise generated by the blade in use.

Experimental and Numerical Study of Impingement on an Airfoil Leading Edge With and Without Showerhead and Gill Film Holes

Abstract: This experimental investigation deals with impingement on the leading-edge of an airfoil with and without showerhead film holes and its effects on heat transfer coefficients on the airfoil nose area as well as the pressure and suction side areas. A comparison between the experimental and numerical results are also made. The tests were run for a range of flow conditions pertinent to common practice and at an elevated range of jet Reynolds numbers (8000–48000). The major conclusions of this study were: a) the presence of showerhead film holes along the leading edge enhances the internal impingement heat transfer coefficients significantly, and b) while the numerical predictions of impingement heat transfer coefficients for the no-showerhead case were in good agreement with the measured values, the case with showerhead flow was underpredicted by as much as 30% indicating a need for a more elaborate turbulence modeling.Copyright © 2005 by ASME

Frame synchronization and initial symbol timing acquisition system and method

Abstract: A robust initial frame detection and symbol synchronization system and methodology are provided. In particular, a first pilot is employed in conjunction with three acquisition stages. In the first stage, an attempt is made to observe the leading edge of the correlation curve associated with the first pilot symbol. In the second stage, a determination is made as to whether a leading edge was detected in the first stage by attempting to observe a flat portion and/or trailing edge of the correlation curve. Furthermore, during this second stage, a frequency loop can be updated to account for frequency offset. The third stage is for observing the trailing edge of the curve if it was not already observed in stage two. Upon detection and confirmation of receipt of the first pilot, a second pilot can subsequently be employed to acquire fine symbol timing.